1. Field of the Invention
This invention relates to the field of flow control in a compressible flow. More specifically, the invention comprises the use of properly placed microjets to create and/or control shock waves in supersonic flow.
2. Description of the Related Art
Shock waves are created whenever an object moves through a compressible fluid at a velocity equal to or greater than the speed of sound in the compressible fluid. Compressible flow in air is of obvious interest, and this application deals primarily with that medium.
Shock wave creation influences many areas of aircraft design. Such waves have a substantial impact on flow over the control surfaces and flow feeding into the powerplant. Effects of the shock waves far away from the aircraft are also a concern. Shock waves striking the ground create a “sonic boom,” which may at best cause annoyance and at worst cause actual damage. In the military context, shock waves may also alert an enemy to the presence of an aircraft which has low visual and radar observability.
The nature of shock waves in the inlets of jet aircraft engines has been thoroughly studied, though the control of such waves has not been thoroughly developed. Those skilled in the art will know that a turbojet can only ingest subsonic air flow. Thus, the duct feeding air into the engine (commonly referred to as an “inlet”) must decelerate the incoming air in a controlled fashion. FIG. 1 shows an idealized prior art inlet 10. It has three main regions spanning the distance between intake 14 and engine compressor 12. These are compression section 18, throat 20, and subsonic diffuser 22.
Supersonic flow passing a discontinuity creates an oblique shock wave—the severity of which depends largely upon the orientation of the discontinuity with respect to the flow (with a more aligned obstruction creating a less severe shock). Oblique shocks 16 are created within compression section 18. These decelerate the flow to a lower Mach number. The oblique shock waves tend to reflect from the contracting walls of the compression section, creating additional deceleration.
During proper operation, the inlet is designed to position normal shock 17 in the vicinity of throat 20. The normal shock is characterized by slightly supersonic flow just upstream and slightly subsonic flow just downstream (typically about Mach 1.3 upstream and about Mach 0.9 downstream). Subsonic diffuser 22 leads from the throat to engine compressor 12. The subsonic flow in this region decelerates as the cavity walls expand. Thus, the flow is further decelerated to well below the speed of sound prior to entering the engine.
The flow within the subsonic diffuser generally obeys Bernoulli's equation in areas away from solid surfaces and where viscous effects are negligible. Such flow is generally predictable (though subsonic compressible flow may often be regarded as counterintuitive). The flow entering and passing through the compression section is more complex, more variable, and more difficult to predict. An effective supersonic inlet design must be able to pass through a wide range of flow states. It must pass from a zero flow state when the engine is started on the ground, through the transonic region, and into supersonic flow.
The shock waves must be created in the correct regions in order for the inlet to perform properly. “Inlet start” refers to the transonic process where the shock waves are established and stabilized. The creation of a stable normal shock in the right position is no small task. In fact, variable incoming and engine flow states may cause “inlet unstart,” a phenomenon wherein the normal shock travels forward into the compression section and may even be ejected out the front of the intake. Such a condition produces drastic flow losses and may flame out the engine.
When the normal shock is in the vicinity of the throat, so that all the subsonic flow lies within the diffuser, the inlet is generally passing the maximum flow. Its operation at this point is said to be critical, which is the desired optimal operating condition. If the normal shock moves aft toward the compressor, the operation is described as supercritical. The flow velocities may be even higher in the supercritical state but the pressure recovery will be reduced, which is generally undesirable.
When the normal shock moves forward toward the intake the operation is described as subcritical. In this state, the inlet is attempting to pass more air than the engine needs. This is generally a dangerous condition since it may cause intake unstart. Bleed flaps or other features upstream of the throat typically remedy such a problem by passing some portion of the intake air around the engine.
Compression inlets have generally been divided into three categories—external compression inlets, internal compression inlets, and mixed compression inlets. External compression inlets locate the oblique shock(s) ahead of the intake. Internal compression inlets locate the oblique shock(s) entirely inside the inlet. Mixed compression inlets have some portion of the oblique shock(s) outside the inlet and some inside the inlet. The spike and cone system on Lockheed's SR-71 Blackbird is an example of an external compression inlet. Most currently produced supersonic aircraft used mixed compression inlets.
The reader will thereby appreciate that the fixed geometry of FIG. 1 will not function in a real aircraft. The fixed geometry shown will only function for a relatively small range of supersonic speeds. It will not serve for subsonic flight, transonic flight, or supersonic flight outside the range. Movable geometry has traditionally been employed to address these concerns.
FIGS. 2 and 3 show two examples of moving inlet geometry, among the many variants known in the art. FIG. 2 shows the diverter/ramp approach. This approach was used on aircraft such as the McDonnell Douglas F-4 Phantom II and the Republic F-105 Thunderchief. Both these aircraft use a pair of inlets lying on either side of the cockpit. FIG. 2(A) corresponds to a “plan view” looking down on one of the aircraft's inlets from above. Boundary layer diverter 26 is a fixed plate mounted a small distance away from fuselage 24. It prevents the turbulent boundary layer air from flowing into the engine(s). FIG. 2(A) shows the subsonic configuration. Adjustable ramp 28 is folded flat against the fuselage, allowing a large intake area. The air flows between the adjustable ramp and lip 34. A compression section, throat and subsonic diffuser are provided before the air reaches engine compressor 12.
FIG. 2(B) shows the geometry set for the supersonic state. Adjustable ramp 28 has been rotated outward—away from the fuselage. First oblique shock 30 is formed by the leading edge of boundary layer diverter 24. Second oblique shock 32 forms at the intersection of the boundary layer diverter and the adjustable ramp. The reader will observe how the position of the adjustable ramp and the shape of outer wall 36 combine to create a converging-diverging path similar to the fixed geometry of FIG. 1. However, since the ramp angle can be changed, the geometry is variable and can function over a broad speed range.
The reader will also observe that both oblique shocks pass to the outside of lip 34. These therefore propagate outside the aircraft. As the shocks in the inlet geometry can be quite strong, the propagation of these waves outside the aircraft contributes to the creation of sonic booms. The state depicted is the low-supersonic region. It has traditionally thought to be desirable to place at least one oblique shock outside the intake so that some decelerated spill flow could pass around the intake. Concerns about sonic booms obviously make this approach less desirable.
FIG. 3 shows another type of variable inlet geometry, such as is found on Grumman's F-14 Tomcat. The view corresponds to an elevation view, looking at one of the aircraft's inlets from the side. Leading lip 33 extends far ahead of trailing lip 35. Two “ramps” are positioned within the inlet—leading ramp 40 and trailing ramp 42. Bleed flap 38 opens and closes to allow some air to bypass the engine—which is located to the right of the geometry shown in the view. FIG. 3(A) shows the subsonic state, in which both ramps are drawn up to the roof of the inlet. A gap between the ramps can be adjusted to allow a certain portion of the air flow to be exhausted through bleed flap 38.
FIG. 3(B) shows the supersonic configuration. The two ramps have been lowered to form throat 20 and subsonic diffuser 22. Leading lip 33 forms first oblique shock 30. Successive oblique shocks are formed by the intersection with leading ramp 40. The throat is formed by the portion of the inlet just behind trailing lip 35 and the repositioned ramps. Some bleed air may still be redirected out through bleed flap 38 in order to properly position the normal shock.
The ramp geometries are typically hydraulically controlled and they can be moved rapidly. Sensors within the inlet—generally pressure sensors and sometimes temperature and flow sensors—monitor the inlet condition. This condition is compared to the engine state in order to properly configure the inlet geometry. All this is done automatically. The reader will note that the oblique shocks in the F-14 inlet are external and propagate beyond the aircraft.
Similar geometry is used on the McDonnell Douglas F-15 Eagle. However, the F-15 incorporates the additional innovative feature of a moving leading lip. The leading lip can be moved forward and downward to reposition the oblique shocks and change the cross section of intake area which is “presented” to the incoming flow. This approach has certain advantages but—as for the Tomcat—oblique shocks are propagated outside the aircraft.
Another factor of great recent interest is radar observability. Most external aircraft features can now be designed to minimize radar reflection. However, the spinning compressor blades produce a strong radar reflection. If any portion of the compressor is visible from the front of the aircraft, the compressor may well be the strongest radar reflection in the entire aircraft. The reader will note that the prior art geometries only partially obscure the compressor. It is therefore desirable to provide a serpentine inlet which will hide the compressor, such as is done in the F-22 Raptor.
A serpentine inlet 10 is shown in FIG. 11. A radar wave entering intake 14 does not directly “see” the plane of engine compressor 12. However, if moving ramps are used in this intake, they may well be visible to radar when deployed. The objective of low observability introduces new flow control concerns, which the present invention may help to address.
As mentioned previously, a mixed compression inlet tends to produce an external oblique shock and a resulting sonic boom. Of course, sonic booms are created by many features outside the engine inlets. FIG. 4 shows the forward portion of an aircraft experiencing supersonic flow. The nose of the aircraft creates a significant shock wave which is a blended combination of bow shock 80 (a normal shock) and oblique shock 16. In addition, every perturbation in the aircraft's surface creates an oblique Mach wave. The joint between the canopy and the fuselage is one good example—“canopy shock” being a thoroughly studied phenomenon. Protuberance 25 on the underside of the fuselage likewise creates an oblique shock.
The mitigation and/or control of shock waves emanating from an aircraft's external structure is a desirable goal. The present invention seeks to create shock waves in desired locations and to control the position and severity of a variety of shock waves. It can be used to control flow and to lessen external effects such as the creation of sonic booms.